Turbofan with motorized rotating inlet guide vane

ABSTRACT

A gas turbine engine according to an exemplary embodiment of this disclosure includes, among other possible things, a fan including a plurality of fan blades rotatable about an axis, a plurality of inlet guide vanes mounted forward of the plurality of fan blades, the plurality of inlet guide vanes selectively rotatable about the axis independent of the plurality of fan blades; and a motor for controlling rotation of the plurality of inlet guide vanes about the axis.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate ahigh-energy exhaust gas flow. The high-energy exhaust gas flow expandsthrough the turbine section to drive the compressor and the fan section.The compressor section typically includes low and high pressurecompressors, and the turbine section includes low and high pressureturbines.

The turbine section drives a fan that provides a substantial portion ofthrust generated by the engine. Guide vanes aft of the fan directairflow into the compressor section to increase compressor efficiency.Airflow into the fan maybe directed to some extent by a nacellestructure surrounding the fan blades, but otherwise is notpre-conditioned.

SUMMARY

A gas turbine engine according to an exemplary embodiment of thisdisclosure includes, among other possible things, a fan including aplurality of fan blades rotatable about an axis, a plurality of inletguide vanes mounted forward of the plurality of fan blades, theplurality of inlet guide vanes selectively rotatable about the axisindependent of the plurality of fan blades; and a motor for controllingrotation of the plurality of inlet guide vanes about the axis.

In a further embodiment of the foregoing gas turbine engine, theplurality of inlet guide vanes are movable from a non-rotating conditionto a rotating condition independent of rotation of the plurality of fanblades.

In another embodiment of any of the foregoing gas turbine engines, theplurality of inlet guide vanes are rotatable at a speed different thanthe plurality of fan blades.

In another embodiment of any of the foregoing gas turbine engines, a fanhub supports rotation of the plurality of fan blades. The plurality ofinlet guide vanes are rotatable, supported by the fan hub for rotationseparate from the fan hub.

In another embodiment of any of the foregoing gas turbine engines, themotor is supported within the fan hub.

In another embodiment of any of the foregoing gas turbine engines, themotor comprises an electric motor.

In another embodiment of any of the foregoing gas turbine engines, theplurality of inlet guide vanes are rotatable about the axis in adirection opposite of rotation of the plurality of fan blades.

In another embodiment of any of the foregoing gas turbine engines, ageared architecture is driven by a turbine section of the gas turbineengine. The geared architecture includes an output driving the pluralityof fan blades at a speed different than the turbine section.

In another embodiment of any of the foregoing gas turbine engines, theplurality of inlet guide vanes comprise an airfoil with a pitch thatdifferent than a pitch of the plurality of fan blades.

A turbofan engine according to an exemplary embodiment of thisdisclosure includes, among other possible things, a fan sectionincluding a plurality of fan blades supported by a fan hub rotatableabout an axis; a compressor section; a combustor in fluid communicationwith the compressor section; a turbine section in fluid communicationwith the combustor; a geared architecture driven by the turbine sectionfor rotating the fan section about the axis; an inlet guide vaneassembly forward of the fan section and rotatable about the axisindependent of the fan section; and a means for controlling rotation ofthe inlet guide vane assembly about the axis independent of the fansection.

In a further embodiment of the foregoing turbofan engine, means forrotating the inlet guide vane assembly is configured to hold the inletguide vane assembly in a fixed position relative to rotation of the fansection.

In a further embodiment of any of the foregoing turbofan engines, theinlet guide vane assembly is rotatable in a direction opposite rotationof the fan section about the axis.

In a further embodiment of any of the foregoing turbofan engines, theinlet guide vane assembly is rotatable at a speed different than a speedof the fan section.

In a further embodiment of any of the foregoing turbofan engines, themeans for rotating the inlet guide vane assembly comprises an electricmotor.

In a further embodiment of any of the foregoing turbofan engines, theinlet guide vane assembly includes a plurality of vanes disposed atpitch different than the plurality of fan blades.

A method of operating a turbofan engine according to an exemplaryembodiment of this disclosure includes, among other possible things,rotating a fan assembly including a plurality of fan blades about anaxis; and imparting a predefined direction of airflow into the fanassembly with a rotatable inlet guide vane assembly disposed forward ofthe fan assembly.

In a further embodiment of the method of operating a turbofan engine,imparting the predefined direction of airflow includes rotating theinlet guide vane in a direction opposite rotation of the fan assemblyabout the axis.

In a further embodiment of any of the foregoing methods of operating aturbofan engine, imparting the predefined direction of airflow includesrotating the inlet guide vane assembly at a speed different than the fanassembly.

In a further embodiment of any of the foregoing methods of operating aturbofan engine, holding the inlet guide vane assembly in a fixedposition relative to rotation of the fan assembly is included.

In a further embodiment of any of the foregoing methods of operating aturbofan engine, the inlet guide vane assembly is driven by an electricmotor supported forward of the fan assembly.

Although the different examples have the specific components shown inthe illustrations, embodiments of this invention are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example gas turbine engine.

FIG. 2 is an enlarged schematic view of a fan section of the example gasturbine engine.

FIG. 3 is a schematic view of the example fan section and inlet guidevane system.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a nacelle18, and also drives air along a core flow path C for compression andcommunication into the combustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith two-spool turbofans as the teachings may be applied to other typesof turbine engines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to a fansection 22 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivefan blades 42 at a lower speed than the low speed spool 30. The highspeed spool 32 includes an outer shaft 50 that interconnects a second(or high) pressure compressor 52 and a second (or high) pressure turbine54. A combustor 56 is arranged in exemplary gas turbine 20 between thehigh pressure compressor 52 and the high pressure turbine 54. Amid-turbine frame 58 of the engine static structure 36 may be arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 58 further supports bearing systems 38in the turbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 58 includes airfoils 60 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor 44 andthe fan blades 42 may be positioned forward or aft of the location ofthe geared architecture 48 or even aft of turbine section 28.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1 and less than about 5:1. Itshould be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine and that thepresent invention is applicable to other gas turbine engines includingdirect drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

The example gas turbine engine includes the fan section 22 thatcomprises in one non-limiting embodiment less than about 26 fan blades42. In another non-limiting embodiment, the fan section 22 includes lessthan about 20 fan blades 42. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about 6 turbine rotorsschematically indicated at 34. In another non-limiting exampleembodiment the low pressure turbine 46 includes about 3 turbine rotors.A ratio between the number of fan blades 42 and the number of lowpressure turbine rotors is between about 3.3 and about 8.6. The examplelow pressure turbine 46 provides the driving power to rotate the fansection 22 and therefore the relationship between the number of turbinerotors 34 in the low pressure turbine 46 and the number of blades 42 inthe fan section 22 disclose an example gas turbine engine 20 withincreased power transfer efficiency.

The disclosed engine 20 includes an inlet guide vane system 64 forwardof the fan section 22 for providing a beneficial pre-swirl and directionto airflow 15 into the fan section 22. The example inlet guide vanesystem 64 includes a plurality of guide vanes 66 that are rotatableabout the engine axis A independent of rotation of the fan blades 42. Amotor 68 is disposed forward of the fan section 22 and enables rotationof the guide vanes 66 independent of rotation of the fan blades 42.

Referring to FIG. 2, with continued reference to FIG. 1, the exampleguide vane system 64 is utilized to precondition and direct airflowentering the fan section 22. The preconditioning or direction of airflowentering the fan section 22 provided by the guide vane system 64enhances efficiency of the fan section 22. Moreover, the guide vanesystem 64 provides for the adjustment of a fan pressure ratio bypreferentially rotating the guide vanes 66 forward of the fan blades 42to preswirl airflow into the fan blades 42.

The fan blades 42 are supported by a fan hub 62 that is rotated, in thisexample, by an output from a geared architecture 48. Accordingly, theexample fan section 22 rotates at a slower speed than the turbine 46driving the fan section 22. Although the disclosed example includes ageared architecture to drive the fan section 22, direct drive engineswould also benefit from this disclosure and is within the contemplationand scope of this disclosure.

The guide vane system 64 includes the motor 68 that rotates a guide vanehub 65 about the axis A. The guide vane hub 65 is rotatable by the motor68 independent of rotation of the fan hub 62. In the disclosedembodiment, the hub 65 is supported by bearings 86 within the rotatingfan hub 62, such that the hub 65 is not subject to rotation with the fanhub 62.

In the disclosed example embodiment, the guide vanes 66 are mounted tothe hub 65 that is mounted on a bearing system 86 that enablesindependent operation and rotation of the inlet guide vanes 66. Aschematic illustration of the example bearing system 86 and hub 65 isshown by way of example. Other configurations that enable separate andindependent operation and rotation of the inlet guide vanes 66 are alsowithin the contemplation and scope of this disclosure.

The motor 68, in this disclosed embodiment, is an electric motor that ispowered through an electric conduit 80 in communication with a motorcontroller 72. The motor controller 72 may be an independent controlleror part of a flight full authority digital engine controller (FADEC),schematically shown at 74. The motor 68 may also provide for holding ofthe guide vanes 66 in a stationary position relative to rotation of thefan blades 42.

The disclosed example engine assembly 20 includes the gearedarchitecture 48 through which the electrical communication conduits 80are directed. The geared architecture 48 provides a convenient paththrough which the electrical conduit 80 can be threaded. Moreover,although a geared fan engine 20 is disclosed by way of example, theexample inlet guide vane system 64 could also be utilized andincorporated into a direct drive engine system.

Referring to FIG. 3, with continued reference to FIG. 2, the guide vanes66 each include a pitch angle 82. The fan blades 42 also include a pitchangle 84. The pitch angle 84 of the fan blades 42 is different than thepitch angle 82 provided by the guide vanes 66. The pitch angle 82provided by the vanes 66 is not intended to generate thrust, but isinstead provided to generate a beneficial and preferential direction toairflow 25 entering the fan section 22. The beneficial airflow 25, inone disclosed example is imparted with a circumferential directedcomponent or pre-swirl that increases fan section propulsive efficiency.

Moreover, the inlet guide vanes 66 do not generate or create a pressureincrease that provide thrust. Instead, the guide vanes 66 rotate in adirection and include the pitch angle 82 that provides for theconditioning and directing of airflow into the fan section 22, such thatthe fan blades 22 can perform more work generating increased thrust.Adjustment and tailoring of a relative speed between the guide vanes 66and the fan blades 42 provide adjustments to the fan pressure ratio.Adjustments to the fan pressure ratio enable tailoring of engineoperation to maximize the thrust generated by the fan section 22 givencurrent environmental and operational conditions.

In operation, inlet airflow 15 enters the nacelle inlet 88 in adirection substantially parallel to the engine axis A. The guide vanes66 impart a preswirl, schematically shown at 25, to the inlet airflow15. The preswirled airflow is communicated to the fan blades 42 in abeneficial manner that increases the efficiency at which the fan bladesgenerate propulsive thrust.

The guide vane 66 may be held in a stationary position or may be rotatedrelative to the fan section 22. In one disclosed example, the guidevanes 66 are rotated in a direction indicated by arrow 90 opposite adirection of the fan blades 42 indicated by arrow 92. Rotation of theguide vanes 66 about the axis A enables the fan section to do more workand achieve a higher pressure ratio for certain operational conditions.Moreover, rotational speed of the guide vane system 64 can be increasedor decreased to adjust the pressure ratio and, thereby, the workperformed by the fan section 22 to tailor operation of the fan section22 to current operational conditions.

Accordingly, the example inlet guide vane system 64 enables conditioningof airflow into the fan section 22 to provide more preferential incomingfan flow to enable tailoring of fan pressure and, thereby, propulsivethrust generated by the fan according to engine environmental operatingconditions.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that this is not intended to be just amaterial specification and that certain modifications would come withinthe scope of this disclosure. For that reason, the following claimsshould be studied to determine the scope and content of this disclosure.

What is claimed is:
 1. A gas turbine engine comprising: a fan includinga plurality of fan blades rotatable about an engine central longitudinalaxis; a geared architecture driven by a turbine section of the gasturbine engine, the geared architecture including an output shaftcoupled to a fan hub supporting the plurality of fan blades for drivingthe plurality of fan blades at a speed slower than the turbine sectionaccording to a gear reduction ratio defined by the geared architecture;a plurality of inlet guide vanes mounted forward of the plurality of fanblades, the plurality of inlet guide vanes selectively rotatable aboutthe engine central longitudinal axis independent of the plurality of fanblades; and a motor coupled to the plurality of inlet guide vanes fordriving rotation of the plurality of inlet guide vanes about the enginecentral longitudinal axis.
 2. The gas turbine engine as recited in claim1, wherein the plurality of inlet guide vanes are movable from anon-rotating condition to a rotating condition independent of rotationof the plurality of fan blades.
 3. The gas turbine engine as recited inclaim 1, wherein the plurality of inlet guide vanes are rotatable at aspeed different than the plurality of fan blades.
 4. The gas turbineengine as recited in claim 1, wherein the plurality of inlet guide vanesare supported by the fan hub for rotation separate from the fan hub. 5.The gas turbine engine as recited in claim 4, wherein the motor issupported within the fan hub.
 6. The gas turbine engine as recited inclaim 1, wherein the motor comprises an electric motor.
 7. The gasturbine engine as recited in claim 1, wherein the plurality of inletguide vanes are rotatable about the axis in a direction opposite ofrotation of the plurality of fan blades.
 8. The gas turbine engine asrecited in claim 1, wherein the plurality of inlet guide vanes comprisean airfoil with a pitch that is different than a pitch of the pluralityof fan blades.
 9. A turbofan engine comprising: a fan section includinga plurality of fan blades supported by a fan hub rotatable about anengine central longitudinal axis; a compressor section; a combustor influid communication with the compressor section; a turbine section influid communication with the combustor; a geared architecture driven bythe turbine section, the geared architecture including an output shaftcoupled to the fan hub for rotating the fan section about the enginecentral longitudinal axis according to a gear reduction ratio defined bythe geared architecture; an inlet guide vane assembly forward of the fansection and rotatable about the engine central longitudinal axisindependent of the fan section and the geared architecture; and a meansfor controlling rotation of the inlet guide vane assembly about theengine central longitudinal axis independent of the fan section.
 10. Theturbofan engine as recited in claim 9, wherein the means for controllingrotation of the inlet guide vane assembly is configured to hold theinlet guide vane assembly in a fixed position relative to rotation ofthe fan section.
 11. The turbofan engine as recited in claim 10, whereinthe inlet guide vane assembly is rotatable in a direction oppositerotation of the fan section about the axis.
 12. The turbofan engine asrecited in claim 9, wherein the inlet guide vane assembly is rotatableat a speed different than a speed of the fan section.
 13. The turbofanengine as recited in claim 9, wherein the means for rotating the inletguide vane assembly comprises an electric motor.
 14. The turbofan engineas recited in claim 9, wherein the inlet guide vane assembly includes aplurality of vanes disposed at pitch different than the plurality of fanblades.
 15. A method of operating a turbofan engine comprising: rotatinga fan assembly including a plurality of fan blades about an enginecentral longitudinal axis with an output shaft driven by a turbinesection through a geared architecture that defines a fixed gearreduction ratio; and imparting a predefined direction of airflow intothe fan assembly with a rotatable inlet guide vane assembly disposedforward of the fan assembly, the rotatable inlet guide vane assemblyrotatable about the engine central longitudinal axis.
 16. The method asrecited in claim 15, wherein imparting the predefined direction ofairflow includes rotating the inlet guide vane in a direction oppositerotation of the fan assembly about the axis.
 17. The method as recitedin claim 15, wherein imparting the predefined direction of airflowincludes rotating the inlet guide vane assembly at a speed differentthan the fan assembly.
 18. The method as recited in claim 15, furtherincluding holding the inlet guide vane assembly in a fixed positionrelative to rotation of the fan assembly.
 19. The method as recited inclaim 15, wherein the inlet guide vane assembly is driven by an electricmotor supported forward of the fan assembly.